Rapid cooling method and apparatus



Oct. 17, 1967 M. c. VAN WANDERHAM ETAL 3,347,057

RAPID COOLING METHOD AND APPARATUS 2 Shecs-Sheet 1 Filed Sept. 27, 1960 FIG.!

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ATTOR NEY 06L 1967 M. c. VAN WANDERHAM ETAL 3,347,057

RAPID COOLING METHOD AND APPARATUS Filed Sept. 27. 1960 2 Sheets-Sheet 2 INVENTORS WARREN W- WORTHLEY MARVIN C- VANWANDERHAM CARL. R- COMOLLI ATTORNEY United States Patent RAPID COQLING METHOD AND APPARATUS Marvin C. Van Wanderham, West Palm Beach, Warren W. Worthley, North Palm Beach, and Carl R. Comolli,

Jupiter, Fla, assignors, by mesne assignments, to the United States of America as represented by the Administrator of the National Aeronautics and Space Administration Filed Sept. 27, 1960, Ser. No. 58,720 3 Claims. (CI. 62-56) This invention relates to cryogenic liquid systems and more particularly to the method and means for accelerating the cooling of the liquid chamber defining walls.

It is an object of this invention to teach method and apparatus for accelerating the cooling of the metal walls of a cryogenic fuel system.'

It is a further object of this invention to provide an insulating layer between the cryogenic fluid and the fluid flow and storage chamber defining walls to prevent film boiling therebetween and to establish nucleate boiling therebetween, thereby accelerating wall cooling and flow stabilization.

It is a further object of this invention to reduce propellant Waste during cool-down of a flight powerplant using cryogenic fuel and thereby saving payload.

Other objects and advantages will be apparent from the specification and claims and from the accompanying drawings which illustrate an embodiment of the invention.

FIG. 1 is a schematic showing of a cryogenic liquid system illustrating our invention.

FIG. 2 is an enlarged and partial cross-sectional showing of a cryogenic liquid vessel Wall, such as a pump wall, having an insulating layer on its inner surface.

FIG. 3 is a graphic illustration of wall cool-down characteristics of a cryogenic fluid system using our teaching and as compared to such a system of conventional construction.

FIG. 4 is a graphic illustration of temperaturev change in the metal Walls of a cryogenic fluid flow system per unit time as the insulating layer thickness is varied to illustrate optimum thickness for a particular wall material, insulator and cryogenic fluid.

In a cryogenic, that is extremely low temperature, liquid containing and flow system such as the rocket fuel supply system shown in FIG. 1, it is essential to cool down or lower the temperature of the cryogenic liquid containing and flow defining walls as rapidly as possible so as to stabilize the flow of cryogenic liquid to such apparatus as rocket thrust chambers and thereby stabilize the operation of such apparatus as rapidly as possible.

In the past, efforts have been made to accelerate this wall cool-down time including the proper selection of cryogenic fluid and Wall metal, providing heat conducting layers therebetween and so forth but none has been acceptable. Quite unexpectedly, applicants while engaged in such work have found that a marked improvement is obtained in cryogenic fluid chamber wall cool-down when an insulating layer is applied to the inner surface of the walls and that optimum efiiciency is obtained when the insulating layer covers the entire wall inner surface. It has been found that the film boiling which was present in the past betwen the wall metal and the cryogenic liquid is prevented from becoming established by the use of this insulating layer and that a nucleate boiling is established in its place. This nucleate boiling allowed better heat transfer than the insulating film boiling between the cryogenic liquid and the walls which occurred when insulation was not used. This film seriously impeded the absorption of heat from the metal walls by the cryogenic liquid. The insulating layer, by the establishment of nucleate boiling, permits some of the cryogenic liquid to be in contact with the insulating layer at all times, even though bubbles are leaving the surface, and thereby accelerates heat transfer between the wall metal and the cryogenic liquid.

We have already tested a highly conductive wall metal of aluminum with insulating materials such as Emralon 310, a tetrafluoroethylene compound and epoxy resin EC 1004 with cryogenic liquids including liquid hydrogen at a temperature of about -420 F. and liquid nitrogen with a temperature of about 320 F.

We call attention to FIG. 3 to illustrate the marked acceleration in wall metal GA AMS 4120 aluminum alloy) cool-down when 20 mils thickness of Emralon 310 insulating layer was used on the inner surface thereof as opposed to an uninsulated application using liquid nitrogen as the cryogenic fluid.

We feel that an optimum insulating layer thickness can be established for each combination of cryogenic fluid, insulating material and wall materials. For example, we have found as illustrated in FIG. 4 that for Emralon 310 as the insulator, the optimum wall thickness is approxi mately .017 inch when the cryogenic fuel is liquid nitrogen and the wall material is aluminum AMS 4120.

To show a workable arrangement for our invention, applicants illustrate a rocket thrust chamber fuel system in FIG. 1 in which a cryogenic fuel and a cryogenic oxidizer are fed to the systems through containers 10 and 12 respectively. Lines 14 and 16 carry the cryogenic propellant to pumps 18 and 20 for controlled distribution through lines 22 and 24 to rocket thrust chamber 26 in which thrust generating combustion occurs. Pumps 18 and 20 are driven through common shaft 28 by turbine 30. Turbine 30 is, in turn, driven by the pro-ducts of combustion from gas generator 32 which receives fuel and oxidizer through lines 34 and 36 at a controlled rate depending upon the setting of throttle valve 38. Comparison control 40 compares the thrust chamber pressure provided to it by pressure sensitive element 42 and the desired thrust chamber pressure provided to it by the pressure sensitive element 44 and controls the flow of fuel and oxidizer to gas generator 32 by establishing the position of throttle valve 38 through positioning mechanism 46 as a function of thrust chamber pressure error. The system shown in FIG. 1 is more fully disclosed in connection with Figs. 8-35 on pages 298 and 299 of Rocket Propulsion Elements by George Sutton. It will be obvious to those skilled in the art that while our invention is readily applicable to the Fig. 1 construction, it is also applicable to innumerable applications including the cryogenic adaptations of embodiments shown in US. Patents Nos. 2,395,- 113, 2,558,483, 2,483,045, and 2,893,202, in which a pressurization is substituted for pumps.

As best shown in FIG. 2, an enlarged cross-sectional showing of a cryogenic liquid vessel wall 50, such as a pump wall, is fabricated to have an insulated layer 52 covering its inner surface and preferably its entire inner surface. An insulating layer such as 52 should preferably cover the inner surfaces of all cryogenic flow defining and storage chambers such as containers 10 and 12, pumps 18 and 20, lines 14, 16, 22 and 24 and so forth.

Those skilled in the art will realize that the aforementioned film boiling and other flow instabilities will cause pump cavitation and that it is therefore highly important to prevent film boiling and to establish cryogenic fluid flow stability as rapidly as possible. As stated above, we have found that the inner surface insulating layer accelerates this metal wall cooling process and prevents film boiling by substituting nucleate boiling therefor.

It is to be understood that the invention is not limited to the specific embodiment herein illustrated and described but may be used in other ways without departure from its spirit as defined by the following claims.

We claim:

1. In a rocket propelled missile booster cryogenic liquid propellant flow'system for uniformly supplying a cryogenic liquid propellant from a propellant storage tank to the combustion chamber in said rocket engine with minimal delay and waste of said fluid in stabilizing said uniform flow, the combination comprising:

(1) conduit means for containing and directing the flow of said cryogenic liquid propellant from said tank to said rocket engine combustion chamber;

(2) a thin substantially uniform layer of thermal insulating plastic material on the internal surfaces of said conduit means in contact with said cryogenic liquid propellant flowing therethrough;

(3) means for producing and maintaining the flow of said cryogenic liquid propellant from said tankto said combustion chamber of said rocket engine;

(4) whereby, the flow of said cryogenic liquid propellant from said tanks to said rocket engine combustion chamber is stabilized.

2. In a rocket propelled missile booster cryogenic liquid propellant flow system for uniformly supplying a cryogenic liquid propellant, from a propellant storage container to the combustion chamber of said rocket engine, the combination comprising:

(1) a cryogenic liquid propellant pump;

(2) a first conduit means for containing and directing the flow of said cryogenic liquid propellant from said storage container to said pump;

(3) a second conduit means for containing and directing the flow of said cryogenic liquid propellant from said pump to said rocket engine combustion chamher; and

(4) a thin uniform layer of thermal insulating plastic material on the internal surfaces of said pump and said first and second conduit means in contact with said cryogenic liquid propellant;

(5) whereby, the flow of said cryogenic liquid propellant from said storage container to said rocket engine combustion chamber is uniform and the velocity and quantity of said flow is controlled by said pump.

3. A method of establishing and maintaining a uniform flow of a cryogenic liquid propellant through a propellant flow system connecting a propellant storage container and a combustion chamber of a rocket engine comprising:

(1) providing a thin uniform layer of thermal insulating plastic material on the internal surfaces of said propellant flow system in contact with said cryogenic liquid propellant when said propellant is flowing through said propellantflow system; and

(2) forcing said cryogenic liquid propellant to flow through said propellant flow system from said propellant storage container to the combustion chamber of said rocket engine;

(3) whereby said cryogenic liquid propellant does not come in direct contact with the interior surfaces of said fuel flow system and said cryogenic liquid pro-. pellant flows through said propellant flow system uniformly with the velocity and quantity of said flow being determined by the means of forcing said cryogenic liquid propellant to flow through said propellant flow system.

References Cited UNITED STATES PATENTS Beckwith 220-9 Moore 220-63 Buckingham 25728 EDWARD J. MICHAEL, Primary Examiner.

ROBERT A. OLEARY, Examiner.

D. R. MATTHEWS, Assistant Examiner.

Morrison 62-54 

1. IN A ROCKET PROPELLED MISSILE BOOSTER CRYOGENIC LIQUID PROPELLANT FLOW SYSTEM FOR UNIFORMLY SUPPLYING A CRYOGENIC LIQUID PROPELLANT FROM A PROPELLANT STORAGE TANK TO THE COMBUSTION CHAMBER IN SAID ROCKET ENGINE WITH MINIMAL DELAY AND WASTE OF SAID FLUID IN STABLIZING SAID UNIFORM FLOW, THE COMBINATION COMPRISING: (1) CONDUIT MEANS FOR CONTAINING AND DIRECTING THE FLOW OF SAID CRYOGENIC LIQUID PROPELLANT FROM SAID TANK TO SAID ROCKET ENGINE COMBUSTION CHAMBER; (2) A THIN SUBSTANTIALLY UNIFORM LAYER OF THERMAL INSULATING PLASTIC MATERIAL ON THE INTERNAL SURFACES OF 